The present invention relates to an integrated panel of interconnected laminates which panel comprises a first laminate and a second laminate, in which the first laminate comprises stacked first metal layers and at least one fibre-reinforced adhesive layer between adjacent metal layers of the first metal layers, in which the second laminate comprises stacked second metal layers and at least one fibre-reinforced adhesive layer between adjacent metal layers of the second metal layers, in which, from the first and second metal layers of the laminates, a metal layer in turn forms an outer metal layer which defines an outer surface of that laminate and said outer surfaces define one and the same continuous and uniformly formed panel surface, in which, from a pair of outer metal layers that comprise a first outer metal layer of the first laminate and a second outer metal layer of the second laminate, the second outer metal layer exhibits a joggled edge section that overlaps an adjacent edge section of the first outer metal layer and that is adhered by means of an adhesive to an adhesion surface of the edge section of the first outer metal layer, which adhesion surface faces away from the panel surface, in which at the location of the transition from the outer surface of the first outer metal layer to the outer surface of the second outer metal layer a filler is located.
Such a panel of fibre metal laminates is known from WO-02/078.950-A1. A fibre metal laminate (FML) has an improved resistance against fatigue, particularly against crack propagation, compared with metal alloys, such as aluminium alloys. Damage resistance is a crucial objective in aerospace engineering. The behaviour of FML in structures such as aircraft structures, for example, FML panels such as fuselage panels from Glare® (glass-fibre composite between aluminium layers), with splices in accordance with the state of the art as known from WO-02/078.950-A1, can however be further improved.
After the manufacture of an FML laminate panel with a splice, a number of additional actions is needed to process the outer edge, so that a smooth outside is achieved that meets the aerodynamic and cosmetic requirements. These actions primarily involve post-processing the outermost filler which is mostly formed as adhesive squeezed out of the splice when manufacturing the panel. This squeeze-out of adhesive is prism-shaped and is freely on the outside of the aircraft. It is disadvantageous that this prism-shaped squeeze-out of adhesive after autoclaving often shows cosmetic imperfections such as small air bubbles, which then have to be post-processed as part of the production process by scratching open, grinding, cleaning, filling with adhesive, hardening for 24 hours, regrinding and painting. That results in a disadvantageous extension of the process for production of such FML laminate panels with one or more splices, in terms of production actions and time.
At the location of the outermost splice overlap, the outer surface of the end region of the outermost metal layer is free, and the thin edge of the outermost metal layer is adhered on the abutting edge to the adjacent outermost prism-shaped squeeze-out of adhesive of the aircraft fuselage panel. This abutting adhesive bonding is unfavourable. In terms of design, adhesive bonds in FML laminates are good at withstanding shear and less good at withstanding tensile forces. The abutting adhesive bonding of the metal edge on the prism-shaped squeeze-out of adhesive, however, is subject to tensile forces, which is not desirable. Due to the repetitive flights of the aircraft, cyclical loads occur in the form of temperature effects (being between −55° C. and 100° C.) and cyclical mechanical load and deformation of the fuselage (cabin pressure and bending moments). Due to these cyclical loads, the outermost squeeze-out of adhesive at the location of the aluminium edge is sensitive to the occurrence of hairline cracks (micro-cracks). If these occur in a painted FML laminate panel, then these would be visible on the outside of the aircraft as cosmetic imperfections in the form of micro-cracks in the paint layer.
Furthermore, in the known splice the outermost metal-metal overlap is provided with a supported adhesive material (provided with a nylon carrier) and the outermost prism of squeeze-out of adhesive comprises a non-supported adhesive material (without carrier), in order to avoid the nylon carrier being able to be drawn from the laminate when processing the prism-shaped squeeze-out of adhesive. Given that supported adhesive is applied in the overlap between the two metal layers, therefore, an extra production action is necessary to deposit the non-supported adhesive for the adjacent outermost prism of squeeze-out of adhesive.
A panel of the type described in the preamble is also found in the book: “Fibre Metal Laminates, an Introduction”, edited by Ad Vlot et al., Kluwer 2001, pages 267-280, Chapter 17: “Detailed design concepts” by O. C. van der Jagt et al.
FML laminates are made as sandwich components of unidirectional glass fibre prepreg, adhesive materials and thin metal plates, for example aluminium plates, assembled onto each other by stacking on a mould. A glass prepreg is a glass fibre mat which is embedded in a matrix material. The metal plates are laid over each other, overlapping in the mould to form splices. After the stacking process, the mould with the product is placed in an autoclave. After that, the product is adhered together in the autoclave under high pressure and high temperature and hardened. As a result of this, a laminate product emerges with a smooth mould side, which also forms the outside of the aircraft.